Cooled pocket in a turbine vane platform

ABSTRACT

An airfoil includes a platform including a pocket, a continuous solid cover plate bonded over the pocket, and a plurality of apertures in the platform in fluid communication with the pocket and adjacent the continuous solid cover plate. A gas turbine engine including the airfoil and a method of cooling an article of a gas turbine engine are also disclosed.

BACKGROUND

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section and a turbine section. Air entering thecompressor section is compressed and delivered into the combustionsection where it is mixed with fuel and ignited to generate a high-speedexhaust gas flow. The high-speed exhaust gas flow expands through theturbine section to drive the compressor and the fan section. Thecompressor section typically includes low and high pressure compressors,and the turbine section includes low and high pressure turbines.

The compressor or turbine sections may include vanes mounted on vaneplatforms. The vanes may require cooling. Some areas that requirecooling cannot be reached by impingement cooling. Therefore, the vaneplatforms may include “cavities” to provide cooling air to the vanes.The cavities may be created in the platforms during casting of theplatforms.

SUMMARY

In a featured embodiment, an airfoil includes a platform including apocket, a continuous solid cover plate bonded over the pocket, and aplurality of apertures in the platform in fluid communication with thepocket and adjacent the continuous solid cover plate.

In another embodiment according to the previous embodiment, the coverplate includes a contour extending into the pocket to define a height ofthe pocket.

In another embodiment according to any of the previous embodiments, thecover plate is bonded over the pocket by being welded to the platform.

In another embodiment according to any of the previous embodiments, atotal cross-section of the plurality of apertures is less thanapproximately 30% of a cross-section of the pocket.

In another embodiment according to any of the previous embodiments, theplurality of apertures comprises a first set of apertures adjacent afirst side of the cover plate and a second set of apertures adjacent asecond side of the cover plate, the second side being opposite the firstside.

In another embodiment according to any of the previous embodiments, thepocket includes a heat-transfer-enhancing surface feature.

In another embodiment according to any of the previous embodiments, afirst side of the platform is adjacent a gas path, and where in thepocket is in a second side opposite the first side.

In another embodiment according to any of the previous embodiments, aheight of the pocket is substantially uniform.

In another featured embodiment, a gas turbine engine includes acompressor section, a turbine section downstream from the compressorsection, and a combustor section in communication with the compressorsection and the turbine section. At least one of the compressor andturbine sections includes a platform including a pocket in a first sideof the platform, an airfoil extending from a second side of the platformopposite the first side, a continuous solid cover plate bonded over thepocket, and a plurality of apertures in the platform adjacent thecontinuous solid cover plate in fluid communication with the pocket.

In another embodiment according the previous embodiment, the pluralityof apertures comprises a first set of apertures adjacent a first side ofthe cover plate and a second set of apertures adjacent a second side ofthe cover plate, the second side being opposite the first side.

In another embodiment according to any of the previous embodiments,cooling air is supplied to the first set of apertures from thecompressor section.

In another embodiment according to any of the previous embodiments, thefirst side of the platform is radially inward from the second side withrespect to a longitudinal axis of the gas turbine engine.

In another embodiment according to any of the previous embodiments, theplatform is a first platform, and further comprising a second platform,the airfoil supported between the first and second platforms.

In another embodiment according to any of the previous embodiments, thefirst platform is radially inward from the second platform with respectto a longitudinal axis of the gas turbine engine.

In another embodiment according to any of the previous embodiments, theairfoil is cantilever-mounted on the platform.

In another embodiment according to any of the previous embodiments, theplatform and airfoil are in the turbine section.

In another featured embodiment according to any of the previousembodiments, a method of cooling an article of a gas turbine engineincludes the steps of supplying cooling air to a pocket in a platformcovered by a continuous solid cover plate via a first set of aperturesadjacent to a first side of the continuous solid cover plate andremoving cooling air from the pocket via a second set of aperturesadjacent to a second side of the continuous solid cover plate, thesecond side being opposite the first side.

Another featured embodiment according to any of the previous embodimentsfurther comprises the step of cooling an article with the cooling airprior to the supplying step.

In another featured embodiment according to any of the previousembodiments, the cooling air is supplied from a compressor section.

In another featured embodiment according to any of the previousembodiments, the article is an airfoil supported on the platform.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of the present disclosure willbecome apparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

FIG. 1 schematically illustrates an example gas turbine engine.

FIG. 2 schematically illustrates a vane assembly for the gas turbineengine of FIG. 1.

FIG. 3 schematically illustrates a cross-section of a platform of thevane assembly of FIG. 2 including a pocket.

FIG. 4A schematically illustrates a cross-section of the platform of thevane assembly of FIG. 2 including an open pocket.

FIG. 4B schematically illustrates a cross-section of the platform of thevane assembly of FIG. 2 including a pocket covered by a cover plate.

FIG. 5 schematically illustrates another cross-section of the platformand airfoil of the vane assembly of FIGS. 4A-4B.

DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures and geared-based engine thatmay or may not have a fan section.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure 36 via several bearing systems38. It should be understood that various bearing systems 38 at variouslocations may alternatively or additionally be provided, and thelocation of bearing systems 38 may be varied as appropriate to theapplication.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five (5:1). Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption ('TSFC')”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tram ° R.)/(518.7°R)]̂0.5. The “Low corrected fan tip speed” as disclosed herein accordingto one non-limiting embodiment is less than about 1150 ft/second.

The high and low pressure compressors 44, 52 and the high and lowpressure turbines 46, 54 may include alternating rows of rotating bladeassemblies 58 and stationary vane assemblies 60. The designs of theblade assemblies 58 and vane assemblies 60 can differ between thecompressor 44, 52 and the turbine 46, 54. A representative vane assembly60 is shown in FIG. 2. The vane assembly 60 of FIG. 2 includes anairfoil 62 supported between first and second platforms 64, 66. However,in another example, the airfoil 62 may be cantilever-mounted on one ofthe first and second platforms 64, 66. In one example, the firstplatform 64 is radially inward from the second platform 66 with respectto the engine axis A (FIG. 1). In the examples of FIGS. 2-5, the firstplatform 64 is shown. However, it is to be understood that the belowdescription is applicable to the second platform 66 as well.

FIG. 3 shows a cross-section of the platform 64 along the section line3-3 in FIG. 2. The platform 64 includes a pocket 72 adjacent a footprintof the airfoil 62. A cover plate 74 covers the pocket 72. The coverplate 74 is continuous and solid. That is, the cover plate is free ofany openings through which gas from the pocket 72 can escape. The coverplate 74 can be bonded to the platform 64 over the pocket 72 so as toseal the pocket 72, for example, by welding, brazing, or the like.Therefore, the pressure inside the pocket 72 when covered by the coverplate 74 can be higher than the pressure outside the pocket 72.

In the example of FIGS. 2-3, the pocket 72 is substantially rectangular.However, in another example, the pocket can be another shape. The pocket72 can be formed in the platform 64 during the process of casting theplatform 64. In another example, the pocket 72 can be formed or bymachining, for example by electrical discharge machining (EDM). In oneexample, the pocket 72 is formed by tooling used to cast the platform64.

Cooling air enters the pocket 72 via one or more inlet apertures 80 inthe platform 64 and exits the pocket 72 via one or more outlet apertures82 in the platform 64. The inlet apertures 80 are adjacent a first side83 a of the pocket 72 and cover plate 74, and the outlet apertures 82are adjacent a second side 83 b of the pocket 72 and cover plate 74opposite the first side 83 a. The cooling air can be cooling air that isused to cool the airfoil 62. The apertures 80, 82 may be, for examplefilm cooling holes. The outlet apertures 82 are open to the exterior ofthe platform 62.

In a modified example shown in FIG. 4A, the pocket 72 is shown open. Inthis further example, the pocket 72 includes one or moreheat-transfer-enhancing surface features 76. In the example of FIG. 4A,the heat-transfer-enhancing surface features 76 are trip strips 76 shownschematically in the internal surface of the pocket 72. The trip strips76 enhance heat transfer from cooling air in the pocket 72. The tripstrips 76 may be ridges that extend obliquely to flow direction throughthe pocket 72, to turbulate the flow. In other examples, theheat-transfer-enhancing surface features 76 can be pedestals or othertypes of features. In further examples, where heat transfer is alreadysufficient, the heat-transfer-enhancing surface features 76 can beexcluded.

In a modified example shown in FIG. 4B, the pocket 72 is covered by thecover plate 74. FIG. 5 shows a cross-section of the platform 64 andairfoil 62 along the section line 5-5 shown in FIG. 4B. In this example,the cover plate 74 is between approximately 12 and 20 mils (0.30 to 0.51mm) thick and includes a contour 78 that extends a depth D into thepocket 72 from a surface of the platform 64. The size of the contour 78defines a height H of the pocket 72. For example, the height H of thepocket 72 is between about 40 and 60 mils (1.02 to 1.52 mm), and issubstantially uniform along a length L of the pocket 72 (FIG. 5).

FIG. 5 shows a cross-section of the the platform 64 and airfoil 62 alongthe line 5-5 (FIG. 4B). The core flow gas path C is adjacent a radiallyoutward side 85 a of the platform 64 with respect to the engine axis A,and the pocket 72 is in on a radially inward side 85 b of the platform64 with respect to the engine axis A. In the example, the inlet aperture80 opens at an inlet on the radially inward side 85 b of the platform 64and extends across a rail 84 of the platform 64. At the opposed outletend, the aperture 80 opens into the pocket 72. In one embodiment, atotal cross-section of apertures 80, 82 opening into the pocket 72 isless than approximately 30% of a cross-section of the pocket 72 in orderto maintain a high pressure in the pocket 72.

In one embodiment, the vane assembly 60 is in one of the low pressureand high pressure turbines 46, 54 and cooling air for impingementcooling the airfoil 62 and for cooling the platform 64 via the pocket 72is supplied from one of the high or low pressure compressors 44, 52,respectively.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of this disclosure. For that reason, the followingclaims should be studied to determine the scope and content of thisdisclosure.

What is claimed is:
 1. An airfoil, comprising: a platform including apocket; a continuous solid cover plate bonded over the pocket; and aplurality of apertures in the platform in fluid communication with thepocket and adjacent the continuous solid cover plate.
 2. The airfoil ofclaim 1, wherein the cover plate includes a contour extending into thepocket to define a height of the pocket.
 3. The airfoil of claim 1,wherein the cover plate is bonded over the pocket by being welded to theplatform.
 4. The airfoil of claim 1, wherein a total cross-section ofthe plurality of apertures is less than approximately 30% of across-section of the pocket.
 5. The airfoil of claim 1, wherein theplurality of apertures comprises a first set of apertures adjacent afirst side of the cover plate and a second set of apertures adjacent asecond side of the cover plate, the second side being opposite the firstside.
 6. The airfoil of claim 1, wherein the pocket includes aheat-transfer-enhancing surface feature.
 7. The airfoil of claim 1,wherein a first side of the platform is adjacent a gas path, and wherein the pocket is in a second side opposite the first side.
 8. Theairfoil of claim 1, wherein a height of the pocket is substantiallyuniform.
 9. A gas turbine engine, comprising: a compressor section; aturbine section downstream from the compressor section; and a combustorsection in communication with the compressor section and the turbinesection; wherein at least one of the compressor and turbine sectionsincludes a platform including a pocket in a first side of the platform,an airfoil extending from a second side of the platform opposite thefirst side, a continuous solid cover plate bonded over the pocket, and aplurality of apertures in the platform adjacent the continuous solidcover plate in fluid communication with the pocket.
 10. The gas turbineengine of claim 9, wherein the plurality of apertures comprises a firstset of apertures adjacent a first side of the cover plate and a secondset of apertures adjacent a second side of the cover plate, the secondside being opposite the first side.
 11. The gas turbine engine of claim10, wherein cooling air is supplied to the first set of apertures fromthe compressor section.
 12. The gas turbine engine of claim 11, whereinthe first side of the platform is radially inward from the second sidewith respect to a longitudinal axis of the gas turbine engine.
 13. Thegas turbine engine of claim 9, wherein the platform is a first platform,and further comprising a second platform, the airfoil supported betweenthe first and second platforms.
 14. The gas turbine engine of claim 13,wherein the first platform is radially inward from the second platformwith respect to a longitudinal axis of the gas turbine engine.
 15. Thegas turbine engine of claim 9, wherein the airfoil is cantilever-mountedon the platform.
 16. The gas turbine engine of claim 9, wherein theplatform and airfoil are in the turbine section.
 17. A method of coolingan article of a gas turbine engine, comprising the steps of: supplyingcooling air to a pocket in a platform covered by a continuous solidcover plate via a first set of apertures adjacent to a first side of thecontinuous solid cover plate; and removing cooling air from the pocketvia a second set of apertures adjacent to a second side of thecontinuous solid cover plate, the second side being opposite the firstside.
 18. The method of claim 17, further comprising the step of coolingan article with the cooling air prior to the supplying step.
 19. Themethod of claim 18, wherein the cooling air is supplied from acompressor section.
 20. The method of claim 18, wherein the article isan airfoil supported on the platform.